The Mach number in the inviscid core of the flow exiting scarfed supersonic nozzles was measured using pitot probes. Nozzle characterization experiments were conducted in a modified section of an obsolete M = 7.3 test section/nozzle assembly on Sandia's Hypersonic Wind Tunnel. By capitalizing on existing hardware, the cost and time required for tunnel modifications were significantly reduced. Repeatability of pitot pressure measurements was excellent, and instrumentation errors were reduced by optimizing the pressure range of the transducers used for each test run. Bias errors in probe position prevented us from performing a successful in situ calibration of probe angle effects using pitot probes placed at an angle to the nozzle centerline. The abrupt throat geometry used in the Baseline and Configuration A and B nozzles modeled the throat geometry of the flight vehicle's spin motor nozzles. Survey data indicates that small (''unmeasurable'') differences in the nozzle throat geometries produced measurable flow asymmetries and differences in the flow fields generated by supposedly identical nozzles. Therefore, data from the Baseline and Configuration A and B nozzles cannot be used for computational fluid dynamics (CFD) code validation. Configuration C and D nozzles replaced the abrupt throat geometry of Baseline and Configuration A and B nozzles with a 0.500-inch streamwise radius of curvature in the throat region. This throat geometry eliminated the flow asymmetries, flow separation in the nozzle throat, and measurable differences between the flow fields from identical nozzles that were observed in Baseline/A/B nozzles. Data from Configuration C and D nozzles can be used for CFD code validation.
Particle image velocimetry (PIV) data have been acquired using three different configurations in the far-field of the interaction of a transverse supersonic jet with a transonic crossflow. The configurations included two-dimensional PIV in the centerline streamwise plane at two overlapping stations, as well as stereoscopic PIV in both the same streamwise plane and the crossplane. The streamwise data show the downstream evolution of the interaction whereas the crossplane data directly reveal its vortex structure. The measurement planes intersect at a common line, allowing a comparison of those mean velocity components and turbulent stresses common to all configurations. All data from the streamwise plane agree to within their estimated uncertainties, but data from the crossplane exhibit reduced velocity and turbulent stress magnitudes by a small but significant degree. Additionally, the vertical positions of the peak velocities are slightly nearer the wall for the crossplane configuration. This comparison suggests that routine methods of uncertainty quantification for data used in the validation of computational models may not fully capture the error sources of an experiment.
A stereoscopic particle image velocimetry (PIV) instrument has been constructed for a transonic wind tunnel to study the interaction created by a supersonic axisymmetric jet exhausting from a flat plate into a subsonic compressible crossflow. Data have been acquired in the crossplane of the interaction at a single station in the farfield, in which the bulk particle motion is aligned with the out-of-plane velocity component. The resulting vector fields distinctly show the strength and location of the induced counter-rotating vortex pair as well as the remnant of the horseshoe vortex that wraps around the jet plume as it first exhausts from the nozzle. Data taken for four different values of the jet-to-freestream dynamic pressure ratio reveal the resulting change in vortex strength, size, and position. Vorticity fields were derived from the in-plane velocity data, but limited convergence of the present small data sets prevented any conclusions about the symmetry of the flowfield. Comparison of the present data is made with two-dimensional PIV data previously acquired in the streamwise plane.
A particle image velocimetry instrument has been constructed for a transonic wind tunnel and applied to study the interaction created by a supersonic axisymmetric jet exhausting from a flat plate into a subsonic compressible crossflow. Data have been acquired in two configurations; one is a two-dimensional measurement on the streamwise plane along the wind tunnel centerline, and the other is a stereoscopic measurement in the crossplane of the interaction. The presence of the induced counter-rotating vortex pair is clearly visible in both data sets. The streamwise-plane data determined the strength and location of the vortices using the vertical velocity component while the crossplane data directly provided a measurement of the vortical motion. A comparison of the vertical velocity component measured using each configuration showed reasonable agreement.
Particle image velocimetry data have been acquired in the far field of the interaction generated by an overexpanded axisymmetric supersonic jet exhausting transversely from a flat plate into a subsonic compressible crossflow. Mean velocity fields were found in the streamwise plane along the flowfield centerline for different values of the crossflow Mach number M{sub {infinity}} and the jet-to-freestream dynamic pressure ratio J. The magnitude of the streamwise velocity deficit and the vertical velocity component both decay with downstream distance and were observed to be greater for larger J while M{sub {infinity}} remained constant. Jet trajectories derived independently using the maxima of each of these two velocity components are not identical, but show increasing jet penetration for larger J. Similarity in the normalized velocity field was found for constant J at two different transonic M{sub {infinity}}, but at two lower M{sub {infinity}} the jet appeared to interact with the wall boundary layer and data did not collapse. The magnitude and width of the peak in the vertical velocity component both increase with J, suggesting that the strength and size of the counter-rotating vortex pair increase and, thus, may have a stronger influence on aerodynamic surfaces despite further jet penetration from the wall.
Despite many decades of jet-in-crossflow experimentation, a distinct lack of data remains for a supersonic jet exhausting into a subsonic compressible crossflow. The present investigation seeks to address this deficiency by examining the flowfield structure of a Mach 3.73 jet injected transversely from a flat plate into a subsonic compressible freestream. The experimental results described herein include the mean surface pressure field as mapped using static pressure taps on the flat plate and an identification of flow features by employing an oil-based surface flow tracer. The possibility of flow separation within the nozzle itself also is addressed using pressure taps along the nozzle interior wall, as is the asymmetry of the separation line due to the variation of the local backpressure around the perimeter of the nozzle orifice resulting from the jet-in-crossflow interaction. Pressure data both on the flat plate and within the nozzle are presented at numerous angles with respect to the crossflow freestream direction to provide a breadth of measurements throughout the interaction region. Since the data are intended for use in validating computational models, attention is paid to providing details regarding the experimental geometry, boundary conditions, flowfield nonuniformities, and uncertainty analyses. Eight different sets of data are provided, covering a range of values of the jet-to-freestream dynamic pressure ratio from 2.8 to 16.9 and a freestream Mach number range of 0.5 to 0.8.
The present document summarizes the experimental efforts of a three-year study funded under the Laboratory Directed Research and Development program of Sandia National Laboratories. The Innovative Diagnostics LDRD project was designed to develop new measurement capabilities to examine the interaction of a propulsive spin jet in a transonic freestream for a model in a wind tunnel. The project motivation was the type of jet/fin interactions commonly occurring during deployment of weapon systems. In particular, the two phenomena of interest were the interaction of the propulsive spin jet with the freestream in the vicinity of the nozzle and the impact of the spin rocket plume and its vortices on the downstream fins. The main thrust of the technical developments was to incorporate small-size, Lagrangian sensors for pressure and roll-rate on a scale model and include data acquisition, transmission, and power circuitry onboard. FY01 was the final year of the three-year LDRD project and the team accomplished much of the project goals including use of micron-scale pressure sensors, an onboard telemetry system for data acquisition and transfer, onboard jet exhaust, and roll-rate measurements. A new wind tunnel model was designed, fabricated, and tested for the program which incorporated the ability to house multiple MEMS-based pressure sensors, interchangeable vehicle fins with pressure instrumentation, an onboard multiple-channel telemetry data package, and a high-pressure jet exhaust simulating a spin rocket motor plume. Experiments were conducted for a variety of MEMS-based pressure sensors to determine performance and sensitivity in order to select pressure transducers for use. The data acquisition and analysis path was most successful by using multiple, 16-channel data processors with telemetry capability to a receiver outside the wind tunnel. The development of the various instrumentation paths led to the fabrication and installation of a new wind tunnel model for baseline non-rotating experiments to validate the durability of the technologies and techniques. The program successfully investigated a wide variety of instrumentation and experimental techniques and ended with basic experiments for a non-rotating model with jet-on with the onboard jets operating and both rotating and non-rotating model conditions.